1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine vanes and the cooling of the leading edge fillet region.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a typical combustion turbine engine, a variety of vortex flows are generated around airfoil elements within the turbine. FIG. 1 is a perspective view of a cut-away of several turbine airfoil portions 1 showing hot combustion fluid flow 3 around the airfoil portions 1. It is known that “horseshoe” vortices, including a pressure side vortex 4, and a suction side vortex 5, are formed when a hot combustion fluid flow 3 collides with the leading edges 6 of the airfoil portions 1. The vortices 4, 5 are formed according to the particular geometry of the airfoil portions 1, and the spacing between the airfoil portions 1 mounted on the platform 2. As the hot combustion fluid flow 3 splits into the pressure side vortex 4 and a suction side vortex 5, the vortices 4, 5 rotate in directions that sweep downward from the respective side of the airfoil portion 1 to the platform 2. On the pressure side 8 of the airfoil portions 1, the pressure side vortex 4 is the predominant vortex, sweeping downward as the pressure side vortex 4 passes by the airfoil portion 1. As shown, the pressure side vortex 4 crosses from the pressure side 8 of the airfoil portion 1 to the suction side 7 of an adjacent airfoil portion 1. In addition, the pressure side vortex 4 increases in strength and size as it crosses from the pressure side 8 to the suction side 7. Upon reaching the suction side 7, the pressure side vortex 4 is substantially stronger than the suction side vortex 5 and is spinning in a rotational direction opposite from the suction side vortex 5. On the suction side 7, the pressure side vortex 4 sweeps up from the platform 2 towards the airfoil portion 1. Consequently, because the pressure side vortex 4 is substantially stronger that the suction side vortex 5, the resultant, or combined flow of the two vortices 4, 5 along the suction side 7 is radially directed to sweep up from the platform 2 towards the airfoil portion 1.
A conventional approach to cooling fluid guide members, such as airfoils in combustion turbines, is to provide cooling fluid, such as high pressure cooling air from the intermediate or last stages of the turbine compressor, to a series of internal flow passages within the airfoil. In this manner, the mass flow of the cooling fluid moving through passages within the airfoil portion provides backside convective cooling to the material exposed to the hot combustion gas. In another cooling technique, film cooling of the exterior of the airfoil can be accomplished by providing a multitude of cooling holes in the airfoil portion to allow cooling fluid to pass from the interior of the airfoil to the exterior surface. The cooling fluid exiting the holes forms a cooling film, thereby insulating the airfoil from the hot combustion gas. While such techniques appear to be effective in cooling the airfoil region, little cooling is provided to the fillet region where the airfoil is joined to a mounting endwall. In a rotor blade, the flow forming surface extending on the sides of the airfoil and root is referred to as a platform. In a stator vane, an inner shroud and an outer shroud that forms the flow surfaces are referred to a endwalls.
The fillet region is important in controlling stresses where the airfoil is joined to the endwall. Although larger fillets can lower stresses at the joint, such as disclosed in U.S. Pat. No. 6,190,128, issued to Fukuno et al on Feb. 29, 2001 and entitled COOLED MOVING BLADE FOR GAS TURBINE the resulting larger mass of material is more difficult to cool through indirect means. Accordingly, prohibitively large amounts of cooling flow may need to be applied to the region of the fillet to provide sufficient cooling. If more cooling flow for film cooling is provided to the airfoil in an attempt to cool the fillet region, a disproportionate amount of cooling fluid may be diverted from the compressor system, reducing the efficiency of the engine and adversely affecting emissions. While forming holes in the fillet to provide film cooling directly to the fillet region would improve cooling in this region, it is not feasible to form holes in the fillet because of the resulting stress concentration that would be created in this highly stressed area.
Backside impingement cooling of the fillet region has been proposed in U.S. Pat. No. 6,398,486. However, this requires additional complexity, such as an impingement plate mounted within the airfoil portion. In addition, the airfoil portion walls in the fillet region are generally thicker, which greatly reduces the effectiveness of backside impingement cooling.
U.S. Pat. No. 6,830,432 B1 issued to Scott et al on Dec. 14, 2004 entitled COOLING OF COMBUSTION TURBINE AIRFOIL FILLETS discloses a row of fillet cooling holes positioned along the airfoil surface just above the fillet extending along the pressure side wall of the airfoil to direct a cooling film over the fillet. FIGS. 4 and 5 show the cooling flows for the Scott et al patent. The Scott et al patent does not disclose any cooling of the fillet in the leading edge region.
As the hot flow core gas enters the turbine with a boundary layer thickness and collides with the leading edge of the vane, the horseshoe vortex separates into a pressure side and suction side downward vortices. Initially, the pressure vortex sweeps downward and flows along the airfoil pressure side forward fillet region first. Then, due to hot flow channel pressure gradient from pressure side to suction side, the pressure side vortex migrates across the hot flow passage and end up at the suction side of the adjacent airfoil. As the pressure side vortex roll across the hot flow channel, the size and strength of the passage vortex becomes larger and stronger. Since the passage vortex is much stronger than the suction side vortex, the suction side vortex flow along the airfoil suction side fillet and acting as a counter vortex for the passage vortex. FIG. 1 shows the vortices formation for a boundary layer entering a turbine airfoil. As a result of these vortices flow phenomena, some of the hot core gas flow from the upper airfoil span is transferred toward close proximity to the end wall and thus creates a high heat transfer coefficient and high gas temperature region at the airfoil fillet region.
As shown in FIG. 1, the resulting forces drive the stagnated flow that occurs along the airfoil leading edge towards the region of lower pressure at the intersection of the airfoil and end wall. This secondary flow flows around the airfoil leading edge fillet and end wall region. This secondary flow then rolls away from the airfoil leading edge and flows upstream along the end wall against the hot core gas flow as seen in FIG. 2. As a result, the stagnated flow forces acting on the hot core gas and radial transfer of hot core gas will flow from the upper airfoil span toward close proximity to the end wall and thus creates a high heat transfer coefficient and high gas temperature region at the intersection location.
Currently, injection of film cooling air at discrete locations along the horseshoe vortex region is used to provide the cooling for this region. However, there are many drawbacks for this type of film blowing injection cooling method. The high film effectiveness level is difficult to establish and maintain in the high turbulent environment and high pressure variation such as horseshoe vortex region. Film cooling is very sensitive to the pressure gradient. The mainstream pressure variation is very high at the horseshoe vortex location. The spacing between the discrete film cooling holes and areas immediately downstream of the spacing are exposed less or provide no film cooling air. Consequently, these areas are more susceptible to thermal degradation and over temperature. As a result of this, spalling of the TBC and cracking of the airfoil substrate will occur.
For the airfoil pressure side fillet region, cooling of the fillet region by means of conventional backside impingement cooling yields inefficient results due to the thickness of the airfoil fillet region. Drilling film cooling holes at the airfoil fillet to provide film cooling produces unacceptable stress by the film cooling holes. An alternative way of cooling the fillet region is by the injection of film cooling air at discrete locations along the airfoil peripheral and end wall into the vortex flow to create a film cooling layer for the fillet region. The film layer migration onto the airfoil fillet region is highly dependent on the secondary flow pressure gradient. For the airfoil pressure side and suction side downstream section, this film injection method provides a viable cooling approach. However, for the fillet region immediately downstream of the airfoil leading edge, where the mainstream or secondary pressure gradient is in the stream-wise direction, injection of film cooling air from the airfoil or end wall surface will not be able to migrate the cooling flow to the fillet region to create a film sub-boundary layer for cooling that particular section of the fillet.
Accordingly, there is a need for improved cooling in the fillet regions of turbine guide members.
It is an object of the present invention to provide for impingement cooling and film cooling of the leading edge fillet region of a turbine vane.